(1) Field of the Invention
The invention relates to a skin-stiffened composite panel with the features of the preamble of claim 1.
(2) Description of Related Art
Aircraft composite structures make use of sandwich designs or skin-stiffened designs. Sandwich designs are typically built of three main constituents: two thin, strong and stiff continuous cover sheets adhesively bonded to each side of a thick core which is considerably weaker and less dense than the cover sheets.
Skin-stiffened designs feature a monolithic skin which is stiffened by longitudinal elements (often denoted as stringers). The stiffeners are either mechanically (riveted), adhesively (bonded or co-bonded) or cohesively (co-cured) attached to the skin, the skin being usually thinner in comparison to the stiffener thickness. The region between the stiffeners is denoted as “bay”.
Monolithic skin-stiffened panels are known to allow for large load bearing capabilities within a post-buckling regime (after buckling of the shear and/or compression loaded skin between the longitudinal stiffeners). The buckled skin excites however large stresses on the main load bearing members and the stiffeners. For metallic panels the operation within a post-buckling regime is not a concern and typical aircraft structures are designed according to the strength capabilities within that regime.
Typically, buckling of the bay skin confined between stiffeners of metallic panels is even allowed below limit load to a specific extent. For composite materials however and especially when having a highly integral design with co-cured, co-bonded or secondary bonded elements, a skin-stiffened design should not be allowed to reach bay skin buckling below limit load, especially for helicopter applications. This is due to the high vibration environment of helicopters considerably increasing the risk of fatigue and damage propagation during a repeated operation beyond the buckling threshold. Skin buckling excites detrimental peeling stresses within the skin-stiffener interface, said peeling stresses being a serious source of delamination. As a result, monolithic skin-stiffened panels require smaller stiffener spacing or/and thicker skins which translates to penalties in terms of the weight efficiency.
Typical composite sandwich designs represent an effective alternative to the skin-stiffened panels enabling a design with large load bearing capabilities before buckling. Sandwich panels are usually not allowed to buckle (or fracture) below ultimate load, since buckling (and local fracture) leads to total structural collapse. Composite sandwich designs are usually (and especially for high load levels) more effective than skin-stiffened designs, as long as post-buckling is not allowed for a skin-stiffened design. However, a sandwich design shows some drawbacks in terms of manufacturing costs, integration, repair, damage tolerance, inspection and operation which make relative their apparent efficiency.
Sandwich designs outstand by their very high bending-stiffness-to-weight and bending-strength-to-weight ratio as well as by a smooth and even shape due to the continuous support of both skins. An additional advantage of sandwich designs is the high thermal insulation in comparison to skin-stiffened designs. The high bending stiffness capabilities of sandwich designs enable large panel dimensions without the need of additional stiffeners. However, sandwich panels are characterized by important drawbacks.
The use of additional adhesive films, minimal thickness requirements (arising from handling, quality and moisture absorption issues), additional hermetic foils, ramp pad-ups and local reinforcements using potting compound reduce the weight efficiency of sandwich panels. The temperature sensitivity related to the adhesive and the core material reduce the operational spectrum. Joining requires heavy inserts and local reinforcements; a later implementation of additional load introduction or attachment points represents a considerable effort. Moreover, the damage resistance and the damage tolerance of sandwich panels with continuous cover sheets are deficient. Especially under fatigue, acoustic, compressive and out-of-plane loading, delaminations caused by an impact tend to propagate all along the interface between the core and the cover sheets leading to important reductions of the panel's residual strength.
The suitability of automatic fiber placement techniques for the manufacturing of typical chamfered sandwich panels is limited. This is due to specific lay-up process issues associated to the ramp geometry and the out-of-plane movement of the roller during lay-up bridging the considerable offset between both facesheets (which is given by the core height, typically ranging between 10 and 20 mm). Typical sandwich panels feature a core made either of cellular (composite or metallic) materials (i.e. honeycomb) or foam materials (i.e. PMI=Polymethacrylimide). Foam materials are usually heavier and less stiff than honeycombs within typical core thickness ranging between 10 and 20 mm.
For lower core thicknesses below 5 mm however, the foam cores can lead to lower overall structural weights than typical honeycombs despite their larger density. This is due to the fact that honeycombs require additional adhesive film layers, whereas foam cores do not necessarily require additional adhesives, especially when using infusion techniques or prepregs with high resin content and corresponding bleeding characteristics.
Contrary to sandwich designs, the panels of skin-stiffened designs feature an easy integration—à posteriori as well—of attachment points, which are typically arranged at the longitudinal stiffeners and show a lower temperature and moisture sensitivity. The quality inspection and repair of said skin-stiffened designs are comparatively easy to process.
However, special attention has to be paid at the interface between the strong stiffeners and the thin skin, especially when adhesively or cohesively coupled, since this interface is prone to delaminate initiating structural damage. This is the reason why composite skin-stiffened designs should not be allowed to reach bay skin buckling during operation, especially when having a high frequency of high load cycles and a high vibration environment as it is the case for helicopter applications. This requirement considerably decreases the weight efficiency of skin-stiffened panels. Skin-stiffened panels are deemed to be more cost-efficient in comparison to sandwich designs, since structural bonding can be fully eliminated, and joining can be accomplished by simple standard means. Moreover, the damage resistance is better, since the skin is thicker and locally more flexible than in the case of a sandwich.
The document DE 10 2008 057 247 B3 discloses a panel having stringers with a hat profile fitted on a skin, where the stringers have an axis in a longitudinal direction, width in a transverse direction and a height as distance of the skin. One of the stringers with the hat profile exhibits a surface curved in a direction of the axis and in the transverse direction. The largest width of the hat profile in the transverse direction of the stringer is set between two adjacent frames. A rib is arranged in a closed area, which is formed between the skin and an internal surface of the stringer.
The document U.S. Pat. No. 7,097,731 B2 discloses a method of manufacturing a hollow section grid-stiffened panel comprising a tool having a surface. The stiffened skin composite panel is preassembled comprising laminating a composite outer skin on the surface, placing a separator outer layer on the composite outer skin, and laminating a composite stiffener on the mandrel, the mandrel being positioned on the separator layer, wherein the separator layer separates the stiffener and the mandrel from the outer composite skin.
The preassembled outer skin composite panel is cured on the tool. The separator layer and mandrel are removed from the preassembled stiffened skin composite panel. The stiffened skin composite panel is reassembled, comprising applying an adhesive between the composite outer skin and the composite stiffener. The reassembled stiffener skin composite panel is cured on the tool to bond the stiffener skin to the outer skin.
The document U.S. Pat. No. 4,452,657 A discloses an “I” beam or web structures of laminated composite material or sheets being stiffened by transverse webs stiffeners having a tubular configuration with flattened end portions. The flattened end is inserted between the laminated web material or sheets and cured therewith to form an integral stiffened web structure.
The document US 2002/0189195 A1 discloses an uncured, thermoset resin sheet reinforced with oriented fibers and is slit to define a desired length for the fibers. A series of the sheets are cut and stacked to form integrated layers of the composite material for a structural panel. The panel has two outer layers that sandwich two shorter layers on each end, and syntactic or foamed resin layer in between. Each of the composite layers is formed from the same materials and by the same process, but may vary in the directional orientation of their fibers.
The uncured panel is assembled into a flexible, substantially flat configuration is heated and formed to the contours of a tool having an undulated surface geometry. The panel is further heated to cure the combined composite and syntactic resins into a series of undulations that permeate each of its layers. In one embodiment, the undulations are in a smooth, sine wave-like pattern that allows the panel to maintain a planar configuration. Although the fibers themselves do not stretch, the short lengths of the fibers enable stretching of the material in the fiber direction so that deformation of the composite is possible in all directions. The panel is stiffened both by spacing its outer layers apart with the syntactic layer and by its undulated surface.
The document US 2009/0309264 A1 discloses a stiffened panel made of a composite with a skin and at least one stiffener having a more or less closed volume. In order for the fibers of the composite to be held in place during fiber deposition and during pressure application while the resin of the composite is being cured, a molding core is placed between the fibers at the position of the more or less closed volume of the stiffener. The molding core includes a flexible bladder filled with a granular solid material, the thermal expansion coefficient of which is close to that of the composite used to produce the stiffened panel. The pressure in the bladder is increased before the composite is cured, so as to compensate for the forces applied for compressing these fibers during production of the panel.
The document US 2011/0311778 A1 discloses a beaded composite panel fabricated using composite plies. An opening is formed in each of plies, and each ply is laid up on a bead feature and drawn down over the bead feature in the area of the opening so as to widen the opening into a gap allowing the ply to conform to the contour of the bead feature. Patches are fabricated and placed on the plies overlying over the openings. The laid-up plies are compacted and cured.